The present invention relates generally to a liquid rocket propulsion system. More particularly, the invention relates to an improved delivery system for conveying the liquid propellants to the thrust chamber and an improved nozzle assembly.
Existing rocket vehicles typically include lower stage booster rocket engines used in conjunction with upper stage sustainer engines. Generally, the booster engines incorporate lower performance, high bulk density propellants, while high-performance liquid propellants tend to be used for the sustainer engines. Hydrogen is the preferred fuel for high-performance liquid rocket engines due to its high thrust per unit of propellant mass flow. However, since the density of hydrogen is very low (only 4.4 lb/ft.sup.3), a relatively large fuel tank is required. In contrast, liquid booster engines often select lower performance fuels such as kerosene because the fuel tank size may be reduced to about one-fourth the size necessary for a corresponding hydrogen tank.
During the past several years NASA and its contractors have been studying relative performance of liquid oxygen/hydrocarbon fuels in booster rocket engines using hydrocarbon fuels such as kerosene (RP-1), propane, and liquid methane. The choice for the best hydrocarbon fuel typically depends upon the selection criteria. A problem associated with the use of hydrocarbon fuels is that they tend to "coat" the inside walls of the coolant channels of the thrust chamber. To overcome this difficulty, a small amount of liquid hydrogen (about 2%) is often assumed to be included with the propellants supplied to cool the thrust chambers. However, such a system requires a tri-propellant booster engine that is complex and potentially unreliable since it requires three propellant tanks and feed systems instead of just two.
As a result of its superior cooling capacity the popularity of liquid hydrogen as a booster rocket fuel has increased in recent years. However, a major drawback of using liquid hydrogen is the required size of its storage tank. Therefore, several attempts have been made to reduce the required size of hydrogen fuel booster tankage. One approach is to operate the liquid oxygen/liquid hydrogen (LOX/LH.sub.2) booster rocket at an oxidizer-rich mixture ratio such as 12:1, as opposed to the typical 6:1 fuel rich mixture ratio. At the higher mixture ratio, the overall bulk density of the propellants are approximately 33 lb/ft.sup.3 as opposed to the 22 lb/ft.sup.3 bulk density of propellants mixed at the lower mixture ratio. The rocket engine thrust per unit engine weight is increased by about the same proportion. A drawback of the higher mixture ratio is that its specific impulse is reduced. However, it remains higher than the specific impulse of liquid oxygen/hydrocarbon booster engines.
For single stage to orbit vehicles of the future, there is a need for a rocket engine that can effectively operate during both the booster and sustainer phases of flight. Combined booster/sustainer rocket engines encounter performance limitations associated with their need to operate both within and outside of the atmosphere. The varying atmospheric pressure at the rocket nozzle exit plane changes the optimum exit-to-throat area ratio from less than 20:1 to more than 100:1 as the vehicle travels from sea level to high altitudes. Concurrently the thrust requirements for the vehicle decrease due to the reduced vehicle mass as a substantial percentage of the propellants are burned, and decreasing gravitational considerations. Therefore, there is a need to provide a booster/sustainer rocket engine having variable thrust capabilities and the capacity to vary the effective nozzle exit-to-throat area ratio.